As aircraft rely more and more on electrical and electronic devices, both for traditional control and communications as well as for flight surface actuation and control on modern fly-by-wire systems, the size and reliability of the electric power generation system must increase to meet these increased utilization requirements. These increasing requirements are typically met through the use of larger generators. However, as the physical size of a generator increases to meet the increased output power requirements, the slower the generator is capable of rotating due to increased stresses and critical speed of the physically larger rotor. Since the speed and size of a generator are inversely related, the actual size of the generator prohibitively increases with the decrease in speed.
This problem is compounded when it is realized that typical electric power generation systems include either generator input speed or output power conversion devices to produce constant frequency ac power at the full aircraft load system rating. This fact is significant because much of the constant frequency power is simply converted to dc power to supply dc loads, or is used by equipment which does not require constant frequency power. Therefore, the extra weight required for the speed or power conversion equipment to produce this quantity of high quality, constant frequency power is to a large degree unneeded. Additionally, since all of the electric power is coupled through the constant frequency bus, any disturbance induced on any downstream distribution bus will be reflected back through this constant frequency bus to all the loads. Additional filtering and scrubbing techniques may be employed to isolate and remove a portion of the disturbance, but this increases the cost and weight of the system.
For these reasons several modern power generation and distribution networks utilize an architecture which segregates the loads into "power quality" and "power type" busses. In this way a physically smaller, higher speed, variable frequency generator may be employed to generate the gross amount of power required on the aircraft. A portion of this variable frequency power is then utilized directly by electrical loads which are not input frequency dependent. Another portion of this power is convened to dc power through a rectifier for use by dc loads, while still another portion of the power is converted to constant frequency power for those loads requiring such high quality power. The result of this architecture is a system which weighs less than a conventional system. Part of the weight reduction is due to the reduced size of the constant frequency power converter, and part is due to the reduced size of the variable speed generator which operates at a higher speed than the conventional constant frequency generator.
This type of system, however, still suffers from the distortion coupling problem of the traditional systems due to the common link of the variable frequency ac bus. An architecture which solves this coupling problem is one which utilizes at least two separate generators, one to supply the ac power and one to supply the dc power. In this way, any distortion induced by, for example, the rectification of the ac output to form a dc output is not coupled back to the variable frequency ac bus. Additionally, loading and faults on one bus do not effect the ability of the other bus to supply the required amount of power. One such system is disclosed in U.S. Pat. No. 4,447,737, which issued on May 8, 1984 to Cronin.
The Cronin system utilizes an induction generator to supply the ac loads and a synchronous permanent magnet generator to supply the dc loads. This system also includes a third synchronous permanent magnet generator which is required to provide excitation to the main induction generator, or, alternatively, to power an ac bus. While this system does not suffer from the coupling problem described above, it does require a spur gear arrangement to drive the separate machines at different speeds. This additional hardware adds weight and cost to the system while reducing overall reliability. Additionally, this system utilizes an induction generator to supply the main ac distribution buses. However, the output power quality of an induction machine is adversely affected during reactive loading conditions. Since typical electrical systems are required to supply power over a range of 0.75 pf lagging to 0.95 pf leading, with motor starting requirements of 0.40 pf lagging, the use of an induction machine to supply main ac power may well be problematic. Also, since many system fault conditions exhibit essentially reactive loading characteristics, an induction generator may not be capable of clearing these faults on the ac bus within acceptable specification limits.
The continuing trend of increased reliance on electrical devices, in addition to requiring that the generating system generate more power, requires that the generating system perform electronic engine starting as well. Many prior systems allowing for electric start of the engines utilize a dedicated starter motor, typically powered by the battery and possibly by an inverter, to generate torque to start the engine. However, this type of arrangement results in a piece of hardware which, although used at the start of the engine, must be carried for the entire flight cycle during which time it is essentially surplusage. The added weight of this additional equipment increases the overall aircraft cost due to increased fuel burn, maintenance, and reliability costs.
The instant invention is directed at overcoming these power generation and engine starting problems known with the prior art systems.